This invention relates to gas turbine engines and, more particularly, to turbine cooling systems for use therein.
Gas turbine engine efficiency is a function of several engine parameters, among them being the temperatures achievable within combustion chambers as well as the amount of air which is extracted from the engine cycle to perform various cooling functions within the engine. The former improves engine cycle efficiency while the latter is detrimental to overall engine performance. Efficiency may also suffer due to excessive amounts of air loss as a result of leakage through malfitting seals. Additionally, the structural integrity of an engine is improved if structural loads are carried by elements of the engine which are not subjected to high temperatures and attendant thermal stresses. The problem is compounded when engine rotational speeds are increased, thus further stressing the structural members. Therefore, the engine designer is continuously faced with a balancing problem of how to increase combustion temperatures and engine rotational speeds in an effort to improve cycle efficiency and yet maintain acceptable stress levels in components subjected to high temperatures and centrifugal loading.
Traditionally, improved methods have been sought to provide improved cooling to the turbine portion of a gas turbine engine, that portion directly downstream of the combustor and through which flow the hot gases of combustion. Accordingly, it has become popular to bleed relatively cool air from the compressor portion of the engine upstream of the combustor and route it rearward through turbine blades provided with circuitous passages in a well-known manner. Cooling is provided by thermal conduction and impingement on internal portions of the turbine blades, with the cooling air typically ejected through a plurality of apertures within the blade body, often so oriented as to provide a cooling film upon the blade airfoil surface. Thus, cooling of the turbine blade is accomplished in an effective manner.
A recently employed concept is that shown in U.S. Pat. No. 3,742,706 -- Klompas, "Dual Flow Cooled Turbine Arrangement for Gas Turbine Engines," assigned to the same assignee as the present invention. Therein, cooling air is bled from the compressor portion through a compressor interstage bleed port and induced radially inwardly through a radial inflow pump. The cooling air passes through an enclosed rotatable annulus to the turbine section where it is drawn into the bore of a hollow turbine disc by means of a centrifugal pump comprising a system of radial vanes or ribs mounted therein. The cooling air is then ejected through the rim of the disc and into the cooperating bases of the turbine blades in the known manner.
However, due to high turbine loading requirements imposed on advanced turbines, the rotor rotational speed is extremely high, often between 25 and 50 percent higher compared to current production engines. This high speed rotation imparts centrifugal forces on all rotor components, these forces being up to twice as large as experienced on conventional engines. Because of this phenomena, a hollow bore entry disc such as advocated by Klompas is additionally required to demonstrate outstanding load-carrying capability while exhibiting the unusual hollow geometrical characteristics needed to deliver cooling air to the turbine blades. It has been discovered that extremely high rotational speeds advocated for such high speed turbines causes the radial vanes (or ribs) to be stressed compressively far in excess of the material capability.
The problem facing the gas turbine designer, therefore, is to provide a highly efficient cooling system for advanced, high speed, high temperature turbines wherein the disc rib compressive stresses are held to acceptable levels within the material limits.